Gas turbine engine with differing effective perceived noise levels at differing reference points and methods for operating gas turbine engine

ABSTRACT

A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular from the turbine that drives the fan. The contribution of the turbine noise emanating from the rear of the engine to the Effective Perceived Noise Level (EPNL) is in the range of from 7 EPNdB and 30 EPNdB lower at a take-off lateral reference point than at an approach reference point.

This is a Continuation-in-Part of application Ser. No. 16/399,123 filedApr. 30, 2019. This application also claims priority to BritishApplication No. 1820943.7 filed Dec. 21, 2018, and British ApplicationNo. 1820939.5 filed Dec. 21, 2018. The entire disclosures of the priorapplications are hereby incorporated by reference herein their entirety.

BACKGROUND

The present disclosure relates to a gas turbine engine having animproved noise signature.

Gas turbine engines are typically optimized to provide high efficiency,because this generally results in lower fuel burn, and thus lowerrunning costs. However, the noise generated by a gas turbine engine usedto power an aircraft is an important factor due to the impact thataircraft noise can have on communities.

In this regard, gas turbine engines generate a significant proportion ofthe noise produced by an aircraft. Regulations define an “EffectivePerceived Noise Level” (EPNL) which is a measure of the impact of thegenerated noise as perceived by the human ear, taking into accountfactors such as frequency, absolute level, tonal components and durationof the noise.

A turbofan gas turbine engine comprises a number of different noisesources. For example, the fan itself is a source of noise, and that fannoise can be separated into two distinct components: a component thatemanates in a forwards direction from the front of the engine; and acomponent that emanates in a rearward direction from the rear of theengine. Further noise sources include (but are not limited to) noisefrom the jet stream exhausted from the engine, noise from the turbine atthe rear of the engine, and noise from the installation of the engine onthe aircraft.

SUMMARY

It is desirable to reduce the perceived noise of a gas turbine engine soas to reduce the impact of the noise on the human ear.

According to an aspect there is provided a gas turbine engine for anaircraft comprising:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and

a gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft, wherein:

the gas turbine is arranged such that during a noise certification testof an aircraft to which the gas turbine is attached,

the contribution of the turbine to the Effective Perceived Noise Level(EPNL) at a take-off lateral reference point is in the range of from7EPNdB and 30 EPNdB lower than the contribution of the turbine to theEPNL at an approach reference noise measurement point, wherein:

the take-off lateral reference point is as defined in Section 3.3.1,a), 1) of the Fifth Edition (July 2008) of Annex 16 (EnvironmentalProtection) to the Convention on International Civil Aviation, Volume 1(Aircraft Noise); and

the approach reference noise measurement point is as defined in Section3.3.1, c) of the Fifth Edition (July 2008) of Annex 16 (EnvironmentalProtection) to the Convention on International Civil Aviation, Volume 1(Aircraft Noise).

According to as aspect there is provided a method of performing a noisecertification test of an aircraft comprising a gas turbine engine,wherein the gas turbine engine comprises:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft and outputs        drive to the fan so as to drive the fan at a lower rotational        speed than the core shaft,

and wherein the noise certification test comprises:

a first test at a take-off lateral reference point as defined in Section3.3.1, a), 1) of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise); and

a second test at an approach reference noise measurement point is asdefined in Section 3.3.1, c) of the Fifth Edition (July 2008) of Annex16 (Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise),

wherein the contribution of the turbine to the Effective Perceived NoiseLevel (EPNL) at the take-off lateral reference point is in the range offrom 7 EPNdB to 30 EPNdB, optionally 10 EPNdB to 30 EPNdB, lower thanthe contribution of the turbine to the EPNL at the approach referencenoise measurement point.

Optionally, in any aspect, the contribution of the turbine to theEffective Perceived Noise Level (EPNL) at the take-off lateral referencepoint may be in a range having a lower bound of any of 7 EPNdB, 10EPNdB, 12EPNdB or 15 EPNdB and an upper bound of any of 30 dB, 28 EPNdBor 25 EPNdB lower than the contribution of the turbine to the EPNL atthe approach reference noise measurement point.

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine, wherein the gas turbine enginecomprises:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft and outputs        drive to the fan so as to drive the fan at a lower rotational        speed than the core shaft, and wherein:

the method comprises taking off from a runway and landing on a runway;and

the contribution of the turbine to the Effective Perceived Noise Level(EPNL) at a take-off lateral reference point, defined as the point on aline parallel to and 450 m from the runway centre line where the EPNLnoise level is a maximum during take-off, is in the range of from 7EPNdB to 30 EPNdB, optionally 10 EPNdB to 30 EPNdB, lower than thecontribution of the turbine to the EPNL at the point on the ground thatis 120 m directly below the path of the aircraft during landing.

According to an aspect there is provided a method of operating a gasturbine engine attached to an aircraft, wherein the gas turbine enginecomprises:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox that receives an input from the core shaft and outputs        drive to the fan so as to drive the fan at a lower rotational        speed than the core shaft, and wherein

the method comprises using the gas turbine engine to provide thrust tothe aircraft for taking off from a runway and landing on a runway; and

the contribution of the turbine to the Effective Perceived Noise Level(EPNL) at a take-off lateral reference point, defined as the point on aline parallel to and 450 m from the runway centre line where the EPNLnoise level is a maximum during take-off, is in the range of from 7EPNdB to 30 EPNdB, optionally 10 EPNdB to 30 EPNdB, lower than thecontribution of the turbine to the EPNL at the point on the ground thatis 120 m directly below the path of the aircraft during landing.

Optionally, in any aspect, the contribution of the turbine to theEffective Perceived Noise Level (EPNL) at the take-off lateral referencepoint may be in a range having a lower bound of any of 7 EPNdB, 10EPNdB, 12EPNdB or 15 EPNdB and an upper bound of any of 30 dB, 28 EPNdBor 25 EPNdB lower than the contribution of the turbine to the EPNL atthe point on the ground that is 120 m directly below the path of theaircraft during landing.

The point on the ground that is 120 m directly below the path of theaircraft during landing may be the point on the ground that is 120 mbelow the lowest point of the fuselage of the aircraft (which may be onthe centreline of the aircraft) at the midpoint between the front andthe rear of the aircraft.

According to an aspect, there is provided a method of operating anaircraft comprising the gas turbine engine as described and/or claimedherein, the method comprising taking off from a runway, wherein themaximum rotational speed of the turbine during take-off is in the rangeof from 5300 rpm to 7000 rpm and/or the maximum rotational speed of theturbine during at the take-off lateral reference point is in the rangeof from 5300 rpm to 7000 rpm.

As referred to herein, including in the claims, the Effective PerceivedNoise Level (EPNL) is as calculated in the conventional manner, asdefined in Appendix 2 of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise). For completeness, the calculationof the EPNL from measured noise data is as defined in Section 4 ofAppendix 2 of the Fifth Edition (July 2008) of Annex 16 (EnvironmentalProtection) to the Convention on International Civil Aviation, Volume 1(Aircraft Noise), from page APP 2-13 to APP 2-21. For completeness, theEPNL is defined at the reference atmospheric conditions provided inSection 3.6.1.5 of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise).

Also as referred to herein, the take-off lateral reference point isdefined as the point on a line parallel to and 450 m from the runwaycentre line where the EPNL is a maximum during take-off, as defined inSection 3.3.1, a), 1) of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise).

Conventionally, the contribution of the turbine to the EPNL at thetake-off lateral reference point would be much closer to—and in somecases the same as, or even higher than—the contribution of the turbineto the EPNL at the approach reference noise measurement point, or at thepoint on the ground that is 120 m directly below the path of theaircraft during landing. Accordingly, the turbine noise at the importanttake-off condition may be reduced, and thus the engine may have a lowernoise impact on communities in the vicinity of the runway.

The present inventors have realised that for a given power of gasturbine engine, the use of a gearbox between the fan and the turbinethat drives the fan enables the turbine noise to be reduced at a highpower condition, such as take-off, because the increased rotationalspeed of the turbine (relative to the fan) allows the frequency of atleast some of the tones generated by the turbine (at least some of whichmay be referred to as fundamental blade passing frequencies) to beincreased, in some cases to levels that are not well perceived (or notat all perceived) by the human ear. Such increased frequencies may alsobe subject to an increase in atmospheric attenuation relative toconventional turbine frequencies. As such, these tones are less wellperceived by the human ear (and possibly at least some tones are not atall perceived by the human ear), and so are given a lower weighting inthe EPNL calculation (even a zero weighting if the frequency is highenough), thereby reducing the contribution of the turbine noise to theEPNL, particularly at conditions where the rotational speed of theturbine is high, such as take-off.

The rotational speed of the turbine that drives the fan via the gearboxmay be greater at the take-off lateral reference point than at theapproach reference noise measurement point. For example, the rotationalspeed of the turbine that drives the fan via the gearbox at the approachreference noise measurement point may be in the range of from 35% to65%, optionally 40% to 60% of the rotational speed of the turbine thatdrives the fan via the gearbox at the take-off lateral reference point.

Specific examples of rotational speed of the turbine that drives the fanvia the gearbox at the take-off lateral reference point are providedelsewhere herein, from which the rotational speed of the turbine thatdrives the fan via the gearbox at the approach reference noisemeasurement point may calculated from the above relationship.

Arrangements of the present disclosure may be particularly beneficialfor fans that are driven via a gearbox. The input to the gearbox may bedirectly from the core shaft, or indirectly from the core shaft, forexample via a spur shaft and/or gear. The core shaft may rigidly connectthe turbine and the compressor, such that the turbine and compressorrotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft. In suchan arrangement, the contribution of the turbine to the EPNL at thetake-off lateral reference point according to any aspect describedand/or claimed herein may be only the contribution of the first turbineto the EPNL at the take-off lateral reference point.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only by the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox is a reduction gearbox (in that the output to the fan is alower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any reduction ratio (defined as the rotational speed ofthe input shaft divided by the rotational speed of the output shaft), asrequired for a particular gas turbine engine to have the relativecontribution of turbine noise and rear fan noise to the EPNL at thetake-off lateral reference point in the ranges described and/or claimedherein. For example, the gear ratio may be greater than 2.5 and/or lessthan 5. By way of more specific example, the gear ratio may be in therange of from 3.2 to 5, or 3.2 to 4.2, or 3.3 to 3.7. By way of furtherexample, the gear ratio may be on the order of or at least 3, 3.1, 3.2,3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2, or between any two ofthe values in this paragraph. In some arrangements, the gear ratio maybe outside these ranges.

The turbine that drives the fan via the gearbox may comprise at leasttwo axially separated rotor stages. For example, the turbine that drivesthe fan via the gearbox may comprise two, three, four, five or greaterthan five axially separated rotor stages. A rotor stage may be part of aturbine stage that also comprises a stator vane stage, which may beaxially separated from the respective rotor stage of the turbine stage.Each rotor stage of the turbine that drives the fan via the gearbox maybe separated from at least one immediately upstream and/or downstreamrotor stage by a row of stator vanes.

The number of turbine blades in the rotor stages of the turbine thatdrives the fan via the gearbox may influence the frequency of at leastsome of the tones generated by the turbine, and thus may assist inallowing the fundamental frequency or frequencies to be moved to a rangewhere they are less well perceived by the human ear (and possibly notperceived at all by the human ear) during take-off, for example at thetake-off lateral reference point

Each and every one of the rotor stages of the turbine that drives thefan via the gearbox may comprise in the range of from 60 to 140 rotorblades, for example in a range having a lower bound of any one of 70,75, 80, 85 or 90, and an upper bound of any one of 140, 130, 120 or 110,for example in the range of from 80 to 140 rotor blades.

The average number of rotor blades in a rotor stage of the turbine thatdrives the fan via the gearbox may be in the range of from 65 to 120rotor blades, for example in a range having a lower bound of any one of65, 70, 75, 80, 85 or 90, and an upper bound of any one of 120, 115, 110or 105, for example in the range of from 85 to 120 rotor blades.

The number of rotor blades in the most axially rearward turbine rotorstage of the turbine that drives the fan via the gearbox may be in therange of from 60 to 120 rotor blades, for example in a range having alower bound of any one of 60, 65, 70, 75, 80, 85 or 90, and an upperbound of any one of 120, 115, 110 or 105, for example 80 to 120 rotorblades.

As noted elsewhere herein, the turbine that drives the fan via thegearbox may comprise at least two axially separated rotor stages. Theturbine that drives the fan via the gearbox may have a rotational speedat the take-off lateral reference point of Wlrp rpm. The minimum numberof rotor blades in any single rotor stage of the turbine that drives thefan via the gearbox may be given by NTURBmin. The diameter of the fanmay be given by fan. A Low Speed System parameter (LSS) may be definedas:

LSS=Wlrp×NTURBmin×ϕfan

The value of the Low Speed System parameter (LSS) may be given by:

1.3×10⁶ m·rpm≤LSS≤2.9×10⁶ m·rpm

The value of the Low Speed System parameter (LSS) may be in a rangehaving a lower bound of any one of 1.3×10⁶ m·rpm, 1.4×10⁶ m·rpm, 1.5×10⁶m·rpm, 1.6×10⁶ m·rpm, 1.7×10⁶ m·rpm, 1.8×10⁶ m·rpm, or 1.9×10⁶ m·rpmand/or an upper bound of any one of 2.9×10⁶ m·rpm, 2.8×10⁶ m·rpm,2.7×10⁶ m·rpm, 2.6×10⁶ m·rpm, 2.5×10⁶ m·rpm, 2.4×10⁶ m·rpm, 2.3×10⁶m·rpm, or 2.2×10⁶ m·rpm.

According to an aspect, which may be independent of or combined withother aspects disclosed herein, there is provided a gas turbine enginefor an aircraft comprising:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and

a gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft, wherein:

the turbine that drives the fan via the gearbox comprises at least twoaxially separated rotor stages and has a rotational speed at thetake-off lateral reference point (as defined elsewhere herein) of Wlrprpm;

the minimum number of rotor blades in any single rotor stage of theturbine that drives the fan via the gearbox is NTURBmin;

the diameter of the fan is fan; and

a low speed system parameter (LSS) is defined as:

LSS=Wlrp×NTURBmin×ϕfan

where:

1.3×10⁶ m·rpm≤LSS≤2.9×10⁶ m·rpm

The value of the Low Speed System parameter (LSS) may be in a rangehaving a lower bound of any one of 1.3×10⁶ m·rpm, 1.4×10⁶ m·rpm, 1.5×10⁶m·rpm, 1.6×10⁶ m·rpm, 1.7×10⁶ m·rpm, 1.8×10⁶ m·rpm, or 1.9×10⁶ m·rpmand/or an upper bound of any one of 2.9×10⁶ m·rpm, 2.8×10⁶ m·rpm,2.7×10⁶ m·rpm, 2.6×10⁶ m·rpm, 2.5×10⁶ m·rpm, 2.4×10⁶ m·rpm, 2.3×10⁶m·rpm, or 2.2×10⁶ m·rpm.

Providing a gas turbine engine with a Low Speed System parameter (LSS)in the ranges described and/or claimed here has been found to result ina gas turbine engine that has high efficiency (for example due inparticular to high propulsive efficiency) and/or high thrust (forexample in the range of from 180 kN to 450 kN), but with acceptably low(and/or lower than conventional) noise levels (for example due toparticularly low turbine noise propagating from the rear of the engine).The rotational speed of the turbine that drives the fan via the gearboxat the take-off lateral reference point may be the same as, or similarto (for example within 5% of), the maximum rotational speed of thatturbine during take-off.

Purely by way of specific example, some gas turbine engines according tothe present disclosure may have a turbine rotational speed at thetake-off lateral reference point (Wlrp) in the range of from 5300 rpm to7000 rpm (for example 5700 rpm to 6500 rpm) and/or a fan diameter in therange of from 320 cm and 400 cm (for example 330 cm and 370 cm) and/or aminimum number of rotor blades in any single rotor stage of the turbinethat drives the fan via the gearbox (NTURBmin) in the range of from 70to 120 (for example 80 to 100).

Purely by way of further specific example, some gas turbine enginesaccording to the present disclosure may have a turbine rotational speedat the take-off lateral reference point (Wlrp) in the range of from 8000rpm to 9500 rpm (for example 8200 rpm to 9200 rpm) and/or a fan diameterin the range of from 220 cm and 290 cm (for example 230 cm and 270 cm)and/or a minimum number of rotor blades in any single rotor stage of theturbine that drives the fan via the gearbox (NTURBmin) in the range offrom 60 to 115 (for example 65 to 115, or 70 to 105).

The total number of turbine blades in the turbine that drives the fanvia the gearbox may be in the range of from 320 and 540, for example inthe range of from 330 to 500, or 340 to 450.

In some arrangements, the relative Mach number at the tip of each fanblade may not exceed 1.09 M at the take-off lateral reference point. Forexample, in some arrangements, the relative Mach number at the tip ofeach fan blade may be in a range having a lower bound of any one of 0.8M, 0.9 M, 1.0 M, 1.01 M or 1.02 M, and an upper bound of any one of 1.05M, 1.06 M, 1.07 M or 1.08 M at the take-off lateral reference point.

Such a relative Mach number at the tip of each fan blade may help toreduce the noise generated by the fan. Accordingly, a gas turbine enginehaving such a relative Mach number at the tip of each fan blade at thetake-off lateral reference point may have both lower fan noise and lowerturbine noise than a conventional engine. As both fan noise and turbinenoise both provide significant contribution to overall engine noise in aconventional engine, such an arrangement may be particularly effectivein reducing overall engine noise. As an additional or alternativebenefit, this, in turn, may reduce the amount of acoustic liner requiredon the intake of the engine, which may facilitate a shorter intake. Forgas turbine engines having a large fan diameter, the intake maycontribute significantly to the aerodynamic drag on the engine duringuse, and so the ability to reduce its extent may be particularlybeneficial.

As used herein, the relative Mach number at the tip of the fan blade maybe defined as the vector sum of the axial Mach number due to the forwardspeed of the engine and the rotational Mach number due to the rotationof the fan blades about the engine axis.

A core nozzle may be defined downstream of the rearmost row of turbineblades in the turbine by a radially inner core nozzle boundary and aradially outer core nozzle boundary. Optionally, at least one of theradially inner core nozzle boundary and the radially outer core nozzleboundary may be provided with a noise attenuation liner. The acousticattenuation liner may comprise acoustic attenuation material whoseprimary, or even sole, function is to attenuate noise. Such a liner maynot be provided to the engine but for its acoustic attenuationproperties. Such a liner may comprise holes which are open to the mainflow (i.e. the flow containing the noise to be attenuated) on one side.Such hopes may open on their other side to cavities. Thus, the noiseattenuation liner may comprise holes that fluidly connect the main flow(e.g. the bypass flow or intake flow) with cavities. The number ofcavities may or may not be equal to the number of holes in such anarrangement.

Optionally, the contribution of the turbine to the EPNL at the take-offlateral reference point is in the range of from 15 dB and 40 dB,optionally 25 dB and 40 dB, lower than the contribution of the fan noiseemanating from the rear of the engine to the EPNL at the take-offlateral reference point.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided downstream of the fan and compressor(s), forexample axially downstream. For example, the combustor may be directlydownstream of (for example at the exit of) the second compressor, wherea second compressor is provided. By way of further example, the flow atthe exit to the combustor may be provided to the inlet of the secondturbine, where a second turbine is provided. The combustor may beprovided upstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 290 cm (for example 230 cm to 270 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 400 cm (for example 330 cm and370 cm) may be in the range of from 1200 rpm to 2000 rpm, for example inthe range of from 1300 rpm to 1800 rpm, for example in the range of from1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 and 20. For example thebypass ratio may be in the range of from 12.5 to 18, for example 13 to17. The bypass ratio may be in an inclusive range bounded by any two ofthe values in the previous sentence (i.e. the values may form upper orlower bounds). The bypass duct may be substantially annular. The bypassduct may be radially outside the core engine. The radially outer surfaceof the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent). Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise conditions.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a schematic showing the measurement of Effective PerceivedNoise Level (EPNL) during take-off;

FIG. 5 is a graph showing an example of how the EPNL varies withdistance during take-off for an example of a gas turbine engine inaccordance with the present disclosure;

FIG. 6 is a graph showing the contribution of turbine noise to the EPNLat the take-off lateral reference point and at an approach condition foran example of a gas turbine engine in accordance with the presentdisclosure;

FIG. 7 is a close-up schematic view of the turbine that drives the fanin an example of a gas turbine engine in accordance with the presentdisclosure; and

FIG. 8 is a diagram illustrating the calculation of fan tip relativeMach number.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Accordingly, the low pressure turbine 19 drivesthe fan 23 via the gearbox 30. Radially outwardly of the planet gears 32and intermeshing therewith is an annulus or ring gear 38 that iscoupled, via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

When in use to power an aircraft, the gas turbine engine 10 generatesnoise. As mentioned elsewhere herein, the gas turbine engine 10according to the present disclosure is arranged to reduce the noiseimpact whilst providing high efficiency.

Take-off is a particularly important flight condition from a noiseperspective, because the engine is typically being operated at a highpower condition, and because the engine is close to the ground, and thuspotentially close to communities. In order to quantify the impact of thegenerated noise as perceived by the human ear, an “Effective PerceivedNoise Level” (EPNL) is defined. The EPNL takes into account factors suchas frequency, absolute level, tonal components and duration of thenoise, and is calculated in the manner defined in Appendix 2 of theFifth Edition (July 2008) of Annex 16 (Environmental Protection) to theConvention on International Civil Aviation, Volume 1 (Aircraft Noise).

A take-off lateral reference point is used in order to quantify theimpact of the generated noise specifically during take-off of anaircraft powered by the gas turbine engine 10, as defined in Section3.3.1, a), 1) of the Fifth Edition (July 2008) of Annex 16(Environmental Protection) to the Convention on International CivilAviation, Volume 1 (Aircraft Noise).

In particular, the take-off lateral reference point is defined as thepoint on a line parallel to and 450 m from the runway centre line wherethe EPNL is a maximum during take-off. This is illustrated in FIG. 4. Inparticular, FIG. 4 shows a series of noise-measurement devices 150, suchas microphones, positioned along a line A on the ground that is 450 mfrom the take-off path (which may be referred to as the runwaycentreline) of an aircraft 100 powered by one or more (for example 2 or4) gas turbine engines 10. Each microphone 150 measures the noise at itsrespective location during take-off, and the measurements are used tocalculate the EPNL at that location. In this way, it is possible todetermine the EPNL along the line A (450 m from the runway centreline,extended forwards along the ground after lift-off).

FIG. 5 shows an example of a graph showing EPNL in dB (EPNdB) along theline A against the distance from the start of take-off (which may bereferred to as distance from release of brake, indicating that it is thedistance from the point at which the aircraft starts its main take-offacceleration at the start of the runway). As illustrated, the EPNL ofthe engine initially increases, and this increase may continue evenafter lift-off (i.e. after the point at which the aircraft loses contactwith the ground), which is labelled as the point “LO” in FIG. 5, purelyby way of example.

At a certain position on the flight path, the EPNL (i.e. the EPNL asmeasured on the ground, along line A in FIG. 4) reaches a maximum, andthen starts to fall. The distance along line A (i.e. the distance alongthe runway centreline) at which this occurs is the take-off lateralreference point (labelled RP in FIG. 5). The EPNL at the take-offlateral reference point RP (labelled RP EPNL in FIG. 5) is the maximumEPNL during take-off.

The take-off period may be considered to last at least as long asnecessary to determine the maximum point (at distance RP) of the EPNLbetween release of brake and top of climb of the aircraft. In practice,this is likely to be within a horizontal distance of 10 km or less ofthe release of brake.

A number of different noise sources contribute to the EPNL, and thus tothe RP EPNL. In a conventional engine, the turbine that drives the fanprovides a significant contribution to the RP EPNL.

However, the present inventors have found that the contribution to theRP EPNL of the turbine 19 that drives the fan 23 via the gearbox 30 canbe significantly reduced by increasing the frequencies of thefundamental tones generated by the turbine to frequencies that are lesswell perceived by the human ear and/or have increased atmosphericattenuation, thereby reducing the perceived noise frequency rating. Assuch, these tones are given a lower weighting in the EPNL calculation(even a zero weighting if the frequency is high enough), therebyreducing the contribution of the turbine noise to the RP EPNL relativeto other noise sources. In FIG. 1, the noise from the turbine 19 thatdrives the fan 23 via the gearbox 30 is illustrated by arrow Q.

FIG. 6 shows the contribution of the noise Q of the turbine 19 thatdrives the fan 23 via the gearbox 30 to the EPNL at approach andtake-off according to an example of the present disclosure. The take-offcondition is the take-off lateral reference point RP, and the approachcondition may be either the approach reference noise measurement point(as defined elsewhere herein) or the point on the ground that is 120 mdirectly below the path of the aircraft during landing. The contributionof the noise Q from the turbine 19 is X EPNdB lower at the take-offcondition than at the approach condition, where X is in the range offrom 7 EPNdB and 30 EPNdB. Purely by way of non-limitative example, thevalue of X for the gas turbine engine 10 having the noise characteristicillustrated in FIG. 6 may be on the order of 25EPNdB.

Accordingly, the gas turbine engine 10 according to the presentdisclosure can be particularly efficient for example having highpropulsive efficiency through having the fan 23 driven via a gearbox 30whilst having reduced noise signature due to the relative reduction innoise (as measured by EPNL) of the turbine 19 that drives the fan 23 viathe gearbox 30 at take-off. Of course, the total engine noise comprisesother noise sources in addition to the turbine noise, such as (by way ofnon-limitative example) jet noise and fan noise (i.e. noise generated bythe fan that emanates from the front and rear of the engine).

It will be appreciated that the individual contributions of thecomponents (such as the noise from the fan 23 that emanates from therear of the engine and the noise from the turbine 19) can be identifiedthrough conventional analysis of the noise measured by the microphones150. For example, each component has a frequency signature that can bepredicted, meaning that noise that is generated in accordance with thepredicted frequency signature can be attributed to that component. Inpractice, the noise that is generated by the fan and emanates from therear of the engine may be distinguished from the noise that is generatedby the fan and emanates from the front of the engine using a sourcelocation technique, such as measuring the phase difference of the noise.In this regard, the noise that is generated by the fan and emanates fromthe rear of the engine is phase-shifted relative to the noise that isgenerated by the fan and emanates from the front of the engine due tothe physical separation of the front and rear of the engine.

FIG. 7 shows the turbine 19 that drives the fan 23 via the gearbox 30 inmore detail for the gas turbine engine 10 according to an example of thepresent disclosure, which may be referred to as the low pressure turbine19. The low pressure turbine 19 comprises four rotor stages 210, 220,230, 240. The low pressure turbine 19 is therefore a four stage turbine19. However, it will be appreciated that the low pressure turbine 19 mayconsist of other numbers of turbine stages, for example three or five.

Each rotor stage 210, 220, 230, 240 comprises rotor blades that extendbetween an inner flow boundary 250 and an outer flow boundary 260. Eachof the rotor stages 210, 220, 230, 240 is connected to the same coreshaft 26 that provides input to the gearbox 30. Accordingly, all of therotor stages 210, 220, 230, 240 rotate at the same rotational speed WIaround the axis 9 in use. In the FIG. 7 example the rotor stages 210,220, 230, 240 each comprise a respective disc 212, 222, 232, 242supporting the rotor blades. However, it will be appreciated that insome arrangements the disc may not be present, such that the blades aresupported on a circumferentially extending disc.

Each rotor stage 210, 220, 230, 240 has an associated stator vane stage214, 224, 234, 244. In use, the stator vane stages do not rotate aroundthe axis 9. Together, a rotor stage 210, 220, 230, 240 and itsassociated stator vane stage 214, 224, 234, 244 may be said to form aturbine stage.

The lowest pressure rotor stage 210 is the most downstream rotor stage.The rotor blades of the lowest pressure rotor stage 210 are longer (i.e.have a greater span) than the rotor blades of the other stages 220, 230,240. Indeed, each rotor stage has blades having a span that is greaterthan the blades of the upstream rotor stages.

The number of rotor blades may have an impact on the frequency of thesound generated by the turbine 19. The rotational speed WI of the lowpressure turbine 19 may also have an effect on the frequency of thesound generated by the turbine 19, and this, in turn, is linked to therotational speed of the fan 23 by the gear ratio of the gearbox 30.

Each rotor stage 210, 220, 230, 240 consists of any desired number ofrotor blades. For example, each and every one of the rotor stages 210,220, 230, 240 of the turbine 19 that drives the fan 23 via the gearbox30 may comprise in the range of from 80 to 140 rotor blades. By way offurther example, the average number of rotor blades in a rotor stage210, 220, 230, 240 of the turbine 19 that drives the fan 23 via thegearbox 30 may be in the range of from 85 to 120 rotor blades. By way offurther example, the number of rotor blades in the most axially rearwardturbine rotor stage 210 of the turbine 19 that drives the fan 23 via thegearbox 30 may be in the range of from 80 to 120 rotor blades.

In one particular, non-limitative example, the first (most upstream)rotor stage 240 and the second rotor stage 230 may each comprise around100 rotor blades, and the third rotor stage 220 and fourth (mostdownstream) rotor stage 210 may each comprise around 90 rotor blades.However, it will be appreciated that this is purely by way of example,and the gas turbine engine 10 in accordance with the present disclosuremay comprise other numbers of turbine blades, for example in the rangesdefined elsewhere herein.

At the take-off lateral reference point, the low pressure turbine 19 hasa rotational speed of Wlrp rpm. In one example, the low pressure turbine19 of gas turbine engine 10 has a rotational speed at the take-offlateral reference point in the range of from 5300 rpm to 7000 rpm. Inthis example, the diameter of the fan 23 (as defined elsewhere herein)may be in the range of from 320 cm to 400 cm. In one specific,non-limitative example, the low pressure turbine 19 of the gas turbineengine 10 has a rotational speed at the take-off lateral reference pointof around 5900 rpm, and a fan diameter of around 340 cm.

In one example, the low pressure turbine 19 of gas turbine engine 10 hasa rotational speed at the take-off lateral reference point in the rangeof from 8000 rpm to 9500 rpm. In this example, the diameter of the fan23 (as defined elsewhere herein) may be in the range of from 220 cm to290 cm. In one specific, non-limitative example, the low pressureturbine 19 of the gas turbine engine 10 has a rotational speed at thetake-off lateral reference point of around 8700 rpm, and a fan diameterof around 240 cm.

A Low Speed System parameter (LSS) may be defined for the gas turbineengine 10 as:

LSS=Wlrp×NTURBmin×ϕfan

where:

Wlrp is the rotational speed at the take-off lateral reference point ofthe turbine 19 that drives the fan 23 via the gearbox 30 (rpm);

NTURBmin is the minimum number of rotor blades in any single rotor stage210, 220, 230, 240 of the turbine 19 that drives the fan 23 via thegearbox 30; and fan is the diameter of the fan (m).

In some arrangements, the Low Speed System parameter (LSS) for the gasturbine engine 10 is in the range:

1.3×10⁶ m·rpm≤LSS≤2.9×10⁶ m·rpm

Purely by way of non-limitative example, the gas turbine engine 10 mayhave a fan diameter of 3.4 m, a minimum number of rotor blades in anysingle rotor stage 210, 220, 230, 240 of 100, and a rotational speed atthe take-off lateral reference point of the low pressure turbine 19 of5900 rpm, giving a Low Speed System parameter (LSS) of around 2.0×10⁶.

Purely by way of further non-limitative example, the gas turbine engine10 may have a fan diameter of 2.4 m, a minimum number of rotor blades inany single rotor stage 210, 220, 230, 240 of 95, and a rotational speedat the take-off lateral reference point of the low pressure turbine 19of 8700 rpm, giving a Low Speed System parameter (LSS) of around2.0×10⁶.

FIG. 8 illustrates a view onto the radially outermost tip of one of thefan blades of the fan 23. In use, the fan blade rotates, such that thetip has a rotational velocity given by the rotational speed of the fanmultiplied by the radius of the tip. The rotational velocity at theleading edge of the tip (i.e. using the radius of the leading edge ofthe tip) can be used to calculate the rotational Mach number at the tip,illustrated by Mn_(rot) in FIG. 8.

The axial Mach number at the leading edge of the tip of the fan blade isillustrated as Mn_(axial) in FIG. 8. In practice (and as used tocalculate the fan tip relative Mach number Mn_(rel) as used herein),this may be approximated by multiplying the average axial Mach numberover the plane that is perpendicular to the axial direction at theleading edge of the tip of the fan blade by 0.9.

The fan tip relative Mach number (Mn_(rel)) is calculated as the vectorsum of the axial Mach number Mn_(axial) and the rotational Mach numberat the tip Mn_(t), i.e. having a magnitude Mnrel=√{square root over(Mnaxial²+Mnrot²)}.

In order to calculate the Mach numbers (Mn_(axial) and Mn_(t)) from thevelocities, the average static temperature over the plane that isperpendicular to the axial direction at the leading edge of the tip ofthe fan blade is used to calculate the speed of sound.

The fan tip relative Mach number (Mn_(rel)) may be in the rangesdescribed and/or claimed herein, for example no greater than 1.09 and/orin the range of from 0.8 M to 1.09 M, optionally 0.9 M to 1.08 M,optionally 1.0 M to 1.07 M at the take-off lateral reference point.

A further example of a feature that may be better optimized for gasturbine engines 10 according to the present disclosure compared withconventional gas turbine engines is the intake region, for example theratio between the intake length L and the fan diameter D. Referring toFIG. 1, the intake length L is defined as the axial distance between theleading edge of the intake and the leading edge of the root of the fanblade, and the diameter D of the fan 23 is defined at the leading edgeof the fan 23. Gas turbine engines 10 according to the presentdisclosure, such as that shown by way of example in FIG. 1, may havevalues of the ratio L/D as defined herein, for example less than orequal to 0.5, for example in the range of from 0.33 to 0.48. This maylead to further advantages, such as installation and/or aerodynamicbenefits, whilst maintaining forward fan noise at an acceptable level.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine for an aircraft, the engine comprising: anengine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; a fan upstream of the enginecore, the fan comprising a plurality of fan blades; and a gearbox thatreceives input from the core shaft and outputs drive to the fan so as todrive the fan at a lower rotational speed than the core shaft, wherein:the gas turbine engine is configured so that contribution of the turbineto Effective Perceived Noise Level (EPNL) at a take-off lateralreference point, defined as a point on a line parallel to and 450 m froma runway centre line where the EPNL is a maximum during take-off, is ina range of from 15EPNdB to 40EPNdB lower than contribution of fan noiseemanating from a rear of the engine to the EPNL at the take-off lateralreference point; the turbine comprises at least two axially separatedrotor stages and has a rotational speed at the take-off lateralreference point of WIrp rpm; a minimum number of rotor blades in anysingle rotor stage of the at least two axially separated rotor stages isNTURBmin; a diameter of the fan is ϕfan; and a low speed systemparameter (LSS) is defined as:LSS=WIrp×NTURBmin×ϕfan where:1.3×10⁶ m·rpm≤LSS≤2.9×10⁶ m·rpm.
 2. The gas turbine engine according toclaim 1, wherein the gas turbine engine is configured so that thecontribution of the turbine to the EPNL at the take-off lateralreference point is in a range of from 20 EPNdB to 40EPNdB lower than thecontribution of the fan noise emanating from the rear of the engine tothe EPNL at the take-off lateral reference point.
 3. The gas turbineengine according to claim 1, wherein the contribution of the turbine tothe EPNL at the take-off lateral reference point is in a range of from25 EPNdB to 40EPNdB lower than the contribution of the fan noiseemanating from the rear of the engine to the EPNL at the take-offlateral reference point.
 4. The gas turbine engine according to claim 1,wherein a relative Mach number at a tip of each of the plurality of fanblades does not exceed 1.09 M when the aircraft is at the take-offlateral reference point.
 5. The gas turbine engine according to claim 4,wherein the relative Mach number at the tip of each of the plurality offan blades is in a range of from 0.8 M to 1.08 M when the aircraft is atthe take-off lateral reference point.
 6. The gas turbine engineaccording to claim 1, wherein: the turbine is a first turbine, thecompressor is a first compressor, and the core shaft is a first coreshaft; the engine core further comprises a second turbine, a secondcompressor, and a second core shaft connecting the second turbine to thesecond compressor; the second turbine, the second compressor, and thesecond core shaft are arranged to rotate at a higher rotational speedthan the first core shaft; and the contribution of the turbine to theEPNL at the take-off lateral reference point is only contribution of thefirst turbine to the EPNL at the take-off lateral reference point. 7.The gas turbine engine according to claim 1, wherein a gear ratio of thegearbox is in a range of from 3.2 to
 5. 8. The gas turbine engineaccording to claim 1, wherein a gear ratio of the gearbox is in a rangeof from 3.2 to 4.2.
 9. The gas turbine engine according to claim 1,wherein each and every one of the at least two axially separated rotorstages comprises in a range of from 60 to 140 rotor blades.
 10. The gasturbine engine according to claim 1, wherein an average number of rotorblades in a said axially separated rotor stage of the turbine is in arange of from 65 to 120 rotor blades.
 11. The gas turbine engineaccording to claim 1, wherein a number of rotor blades in a most axiallyrearward of the at least two axially separated rotor stages is in arange of from 60 to 120 rotor blades.
 12. (canceled)
 13. The gas turbineengine according to claim 1, wherein a total number of turbine blades inthe turbine is in a range of from 320 and
 540. 14. The gas turbineengine according to claim 1, wherein the diameter of the fan is in arange of from 320 cm and 400 cm.
 15. The gas turbine engine according toclaim 1, wherein a bypass ratio of the gas turbine engine at cruiseconditions is in a range of from 12 to
 18. 16. A method of operating anaircraft comprising the gas turbine engine according to claim 1, themethod comprising taking off from a runway, wherein a maximum rotationalspeed of the turbine during take-off is in a range of from 5300 rpm to7000 rpm.
 17. A method of operating a gas turbine engine attached to anaircraft, wherein: the gas turbine engine comprises: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan upstream of the engine core, the fancomprising a plurality of fan blades; and a gearbox that receives inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft; the turbine comprisesat least two axially separated rotor stages and has a rotational speedat a take-off lateral reference point, defined as a point on a lineparallel to and 450 m from a runway centre line where EffectivePerceived Noise Level (EPNL) is a maximum during take-off, of WIrp rpm;a minimum number of rotor blades in any single rotor stage of the atleast two axially separated rotor stages is NTURBmin; a diameter of thefan is ϕfan; a low speed system parameter (LSS) is defined as:LSS=WIrp×NTURBmin×ϕfan where:1.3×10⁶ m·rpm≤LSS≤2.9×10⁶ m·rpm; and the method comprises using the gasturbine engine to provide thrust to the aircraft for taking off from arunway so that contribution of the turbine to the EPNL at the take-offlateral reference point is in a range of from 15EPNdB to 40EPNdB lowerthan contribution of fan noise emanating from a rear of the engine tothe EPNL at the take-off lateral reference point.
 18. A method ofoperating an aircraft comprising a gas turbine engine, wherein: the gasturbine engine comprises: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan upstream of the engine core, the fan comprising a plurality of fanblades; and a gearbox that receives input from the core shaft andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the core shaft; the turbine comprises at least two axiallyseparated rotor stages and has a rotational speed at a take-off lateralreference point, defined as a point on a line parallel to and 450 m froma runway centre line where Effective Perceived Noise Level (EPNL) is amaximum during take-off, of WIrp rpm; a minimum number of rotor bladesin any single rotor stage of the at least two axially separated rotorstages is NTURBmin; a diameter of the fan is ϕfan; a low speed systemparameter (LSS) is defined as:LSS=WIrp×NTURBmin×ϕfan where:1.3×10⁶ m·rpm≤LSS≤2.9×10⁶ m·rpm; and the method comprises taking offfrom a runway so that contribution of the turbine to the EPNL at thetake-off lateral reference point is in a range of from 15EPNdB to40EPNdB lower than contribution of fan noise emanating from a rear ofthe engine to the EPNL at the take-off lateral reference point.
 19. Themethod according to claim 17, wherein1.6×10⁶ m·rpm≤LSS≤2.8×10⁶ m·rpm.
 20. The gas turbine engine according toclaim 3, wherein the contribution of the turbine to the EPNL at thetake-off lateral reference point is in a range of from 25 EPNdB to 35EPNdB lower than the contribution of the fan noise emanating from therear of the engine to the EPNL at the take-off lateral reference point.21. The gas turbine engine according to claim 8, wherein the gear ratioof the gearbox is in a range of from 3.3 to 3.7.
 22. The gas turbineengine according to claim 9, wherein each and every one of the at leasttwo axially separated rotor stages comprises in a range of from 80 to140 rotor blades.
 23. The gas turbine engine according to claim 10,wherein an average number of rotor blades in a said axially separatedrotor stage of the turbine is in a range of from 85 to 120 rotor blades.24. The gas turbine engine according to claim 11, wherein the number ofrotor blades in the most axially rearward of the at least two axiallyseparated rotor stages is in a range of from 80 to 120 rotor blades. 25.The gas turbine engine according to claim 1, wherein1.6×10⁶ m·rpm≤LSS≤2.8×10⁶ m·rpm.
 26. The gas turbine engine according toclaim 15, wherein the bypass ratio of the gas turbine engine at cruiseconditions is in a range of from 13.0 to 18.0.
 27. The method accordingto claim 18, wherein1.6×10⁶ m·rpm≤LSS≤2.8×10⁶ m·rpm.